The harvesting of a large cross-section area of solar energy from a smaller area of solar cells by concentration is a well-recognized art. Known techniques include using lenses such as fresnel lenses to focus the energy onto the cells, and large mirror arrangements to reflect the energy to the cells. However effective these known devices may be in directing energy from a larger area onto a smaller area, they bring with them many problems of practical concern when used for spacecraft.
Land-based apparatus to capture solar energy involves no special problems. Weight, rigidity, variability of environmental conditions such as temperature, and reliability are of lesser concern or may be of concern at all. Most of them can be minimized or corrected by over-design of the apparatus. It is built in place and stays there. Temperature variability is relatively minor. Weight is no concern. Neither is perfect reliability, because within reason the apparatus is accessible and readily repaired. Structural efficiency is really not an issue. A ground based device may simply be made as heavy and strong as desired, with a generous allowance for safety.
Such is not the circumstance for spacecraft. Weight is a primary consideration not only because of the cost per pound to launch the apparatus, but because weight of one part of a spacecraft will necessarily require a reduction of weight elsewhere due to the ultimate limitation on the total launch weight of the entire craft.
Reliability is also a prime concern. Spacecraft are one-way vehicles. Once in space they remain there during their useful life, and except in a few extraordinary situations such as the Hubble Telescope, they will never be approached after launch. The failure of apparatus such as a solar array dooms all or a large part of the intended life and function of the entire craft.
Rigidity in the sense of maintenance of shape under varying conditions is made complicated by the extreme variations in temperature as the apparatus enters and leaves the shadow of the earth. While in the shadow, temperatures as low as -180 degrees C are endured. While out of the shadow and exposed directly to the sun, temperatures as high as 110 degrees C are endured. When the solar panels transition between light and shadow, the change in temperature of the apparatus occurs in only a few minutes, and does not occur uniformly throughout. This results in a reaction known as "thermal snap" in which the distortions that result from rapid temperature change cause a quick bending distortion that shudders the spacecraft and can damage the wing.
As the wing temperatures change over the sunlit portion of the orbit, distortions of the structure can cause the concentrator optics to malfunction. Even more, high temperatures are the enemy of solar cells. It is important to mount the cells in an arrangement such that the energy received by them does not heat the entire array of cells to an unacceptable temperature. This is further complicated if large reflecting areas are involved where there may be localized higher temperatures due to distortions of the reflector.
This is a daunting array of requirements. Over the decades there has been a long succession of solar arrays produced and launched. Many have been successful, but a disheartening proportion of them have failed partially or totally, causing the loss of very costly space vehicles, or a major reduction in their useful life.
It is an object of this invention to provide a lightweight, structurally integral solar array that is readily packed for launch. Upon being opened in space it will itself erect to an operational configuration, without requiring energy or exertion from another source.
It is another object to provide a structure which is forgiving of both global and localized temperature variations, and whose shape is inherently biased toward the optimum.